Cryogenic fuel system and method for delivering fuel in an aircraft

ABSTRACT

A cryogenic fuel system for an aircraft having a turbine engine with a compressor section and a combustion chamber, including a tank for storing cryogenic fuel, a supply line operably coupling the tank to the combustion chamber and a pump coupling the tank to the supply line to pump the cryogenic fuel at high pressure through the supply line where the pump is operably coupled to the compressor such that operation of the turbine engine drives the pump and a method for delivering fuel in a fuel system to a turbine engine.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of U.S. Provisional PatentApplication No. 61/747,169, filed on Dec. 28, 2012, which isincorporated herein in its entirety.

BACKGROUND OF THE INVENTION

The technology described herein relates generally to aircraft systems,and more specifically to aircraft systems using dual fuels in anaviation gas turbine engine and a method of operating same.

Certain cryogenic fuels such as liquefied natural gas (LNG) may becheaper than conventional jet fuels. Current approaches to cooling inconventional gas turbine applications use compressed air or conventionalliquid fuel. Use of compressor air for cooling may lower efficiency ofthe engine system.

Accordingly, it would be desirable to have aircraft systems using dualfuels in an aviation gas turbine engine. It would be desirable to haveaircraft systems that can be propelled by aviation gas turbine enginesthat can be operated using conventional jet fuel and/or cheapercryogenic fuels such as liquefied natural gas (LNG). It would bedesirable to have more efficient cooling in aviation gas turbinecomponents and systems. It would be desirable to have improvedefficiency and lower Specific Fuel Consumption in the engine to lowerthe operating costs. It is desirable to have aviation gas turbineengines using dual fuels that may reduce environmental impact with lowergreenhouse gases (CO2), oxides of nitrogen—NOx, carbon monoxide—CO,unburned hydrocarbons and smoke.

BRIEF DESCRIPTION OF EMBODIMENTS OF THE INVENTION

In one aspect, an embodiment of the invention relates to a cryogenicfuel system for an aircraft having a turbine engine with a compressorsection and a combustion chamber, including a tank for storing cryogenicfuel, a supply line operably coupling the tank to the combustion chamberand a pump coupling the tank to the supply line to pump the cryogenicfuel at high pressure through the supply line where the pump is operablycoupled to the compressor such that operation of the turbine enginedrives the pump.

In another aspect, an embodiment of the invention relates to a methodfor delivering fuel in a fuel system to a turbine engine having acombustion section and a compressor section, including pressurizing thefuel provided to the combustion section of the turbine engine utilizinga turbopump driven by compressor discharge pressure from the turbineengine.

BRIEF DESCRIPTION OF THE DRAWINGS

The technology described herein may be best understood by reference tothe following description taken in conjunction with the accompanyingdrawing figures in which:

FIG. 1 is an isometric view of an exemplary aircraft system having adual fuel propulsion system;

FIG. 2 illustrates an exemplary pressurization system embodiment;

FIG. 3 is an exemplary fuel delivery/distribution system;

FIG. 3a is an exemplary operating path in a schematic pressure-enthalpychart of an exemplary cryogenic fuel;

FIG. 4 is a schematic figure showing exemplary arrangement of a fueltank and exemplary boil off usage;

FIG. 5 is a schematic cross-sectional view of an exemplary dual fuelaircraft gas turbine engine having a fuel delivery and control system;

FIG. 6 is a schematic cross-sectional view of a portion of an exemplarydual fuel aircraft gas turbine engine showing a schematic heatexchanger;

FIG. 7a is a schematic view of an exemplary direct heat exchanger;

FIG. 7b is a schematic view of an exemplary indirect heat exchanger;

FIG. 7c is a schematic view of another exemplary indirect heatexchanger; and

FIG. 8 is a schematic plot of an exemplary flight mission profile forthe aircraft system.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

Referring to the drawings herein, identical reference numerals denotethe same elements throughout the various views.

FIG. 1 shows an aircraft system 5 according to an exemplary embodimentof the present invention. The exemplary aircraft system 5 has a fuselage6 and wings 7 attached to the fuselage. The aircraft system 5 has apropulsion system 100 that produces the propulsive thrust required topropel the aircraft system in flight. Although the propulsion system 100is shown attached to the wing 7 in FIG. 1, in other embodiments it maybe coupled to other parts of the aircraft system 5, such as, forexample, the tail portion 16.

The exemplary aircraft system 5 has a fuel storage system 10 for storingone or more types of fuels that are used in the propulsion system 100.The exemplary aircraft system 5 shown in FIG. 1 uses two types of fuels,as explained further below herein. Accordingly, the exemplary aircraftsystem 5 comprises a first fuel tank 21 capable of storing a first fuel11 and a second fuel tank 22 capable of storing a second fuel 12. In theexemplary aircraft system 5 shown in FIG. 1, at least a portion of thefirst fuel tank 21 is located in a wing 7 of the aircraft system 5. Inone exemplary embodiment, shown in FIG. 1, the second fuel tank 22 islocated in the fuselage 6 of the aircraft system near the location wherethe wings are coupled to the fuselage. In alternative embodiments, thesecond fuel tank 22 may be located at other suitable locations in thefuselage 6 or the wing 7. In other embodiments, the aircraft system 5may comprise an optional third fuel tank 123 capable of storing thesecond fuel 12. The optional third fuel tank 123 may be located in anaft portion of the fuselage of the aircraft system, such as for exampleshown schematically in FIG. 1.

As further described later herein, the propulsion system 100 shown inFIG. 1 is a dual fuel propulsion system that is capable of generatingpropulsive thrust by using the first fuel 11 or the second fuel 12 orusing both first fuel 11 and the second fuel 12. The exemplary dual fuelpropulsion system 100 comprises a gas turbine engine 101 capable ofgenerating a propulsive thrust selectively using the first fuel 11, orthe second fuel 21, or using both the first fuel and the second fuel atselected proportions. The first fuel may be a conventional liquid fuelsuch as a kerosene based jet fuel such as known in the art as Jet-A,JP-8, or JP-5 or other known types or grades. In the exemplaryembodiments described herein, the second fuel 12 is a cryogenic fuelthat is stored at very low temperatures. In one embodiment describedherein, the cryogenic second fuel 12 is Liquefied Natural Gas(alternatively referred to herein as “LNG”). The cryogenic second fuel12 is stored in the fuel tank at a low temperature. For example, the LNGis stored in the second fuel tank 22 at about −265° F. at an absolutepressure of about 15 psia. The fuel tanks may be made from knownmaterials such as titanium, Inconel, aluminum or composite materials.

The exemplary aircraft system 5 shown in FIG. 1 comprises a fueldelivery system 50 capable of delivering a fuel from the fuel storagesystem 10 to the propulsion system 100. Known fuel delivery systems maybe used for delivering the conventional liquid fuel, such as the firstfuel 11. In the exemplary embodiments described herein, and shown inFIGS. 1 and 3, the fuel delivery system 50 is configured to deliver acryogenic liquid fuel, such as, for example, LNG, to the propulsionsystem 100 through conduits 54 that transport the cryogenic fuel. Inorder to substantially maintain a liquid state of the cryogenic fuelduring delivery, at least a portion of the conduit 54 of the fueldelivery system 50 is insulated and configured for transporting apressurized cryogenic liquid fuel. In some exemplary embodiments, atleast a portion of the conduit 54 has a double wall construction. Theconduits may be made from known materials such as titanium, Inconel,aluminum or composite materials.

The exemplary embodiment of the aircraft system 5 shown in FIG. 1further includes a fuel cell system 400, comprising a fuel cell capableof producing electrical power using at least one of the first fuel 11 orthe second fuel 12. The fuel delivery system 50 is capable of deliveringa fuel from the fuel storage system 10 to the fuel cell system 400. Inone exemplary embodiment, the fuel cell system 400 generates power usinga portion of a cryogenic fuel 12 used by a dual fuel propulsion system100.

One of the challenges in utilizing cryogenic fuels in gas turbineapplications is that the fuel needs to be injected into the combustionsection of the jet engine 101 at a pressure higher than the pressure ofair coming into the combustor or combustion section from the compressorsection of the gas turbine. The pressure delta between the fuel and theair needs to be high enough to overcome the pressure drop across allcomponents in the fuel line from the fuel tank to the fuel nozzles, aswell as across the fuel nozzles to ensure proper stoichiometric fuel toair ratios as well as to avoid back flow. Cryogenic fuel may bepressurized in a variety of manners including: 1) heating the fuelinside the tank and allowing it to boil off within the tank, whichpressurizes the fuel within the tank; 2) utilizing an electricallydriven pump to pressurize the fuel downstream of the tank; 3) utilizinga turbopump, powered by compressor discharge pressure to pressurize thefuel. Utilizing a turbopump poses significant advantage in reducingcomplexity, weight, and external power requirements.

FIG. 2 illustrates a cryogenic fuel system 500 for an aircraft 5 havinga turbine engine 506 with a compressor section and a combustion chamber.The cryogenic fuel system includes a tank 515 for storing cryogenicfuel, a supply line 514 operably coupling the tank 515 to the combustionchamber and a pump 502 coupling the tank 515 to the supply line 514 topump the cryogenic fuel at high pressure through the supply line 514.The pump 502 is operably coupled to the compressor, as indicated at 504,such that operation of the jet engine drives the pump 502.

More specifically, the cryogenic fuel system 500 may utilize turbopumptechnology to pressurize cryogenic fuels in gas turbine applications. Asillustrated, it is contemplated that the turbine engine 506 may beconfigured to be fueled by a first fuel and a second fuel and mayinclude a combustor section and a compressor section. For example, afirst fuel system 530 for delivering a first fuel from a first fuel tank532 to the turbine engine 506 has been illustrated along with the secondfuel system or the cryogenic fuel system 500.

The pump 502 may pressurize the second fuel to a pressure higher thanthe pressure of air coming into the combustion section of the turbineengine 506 from the compressor section of the turbine engine 506. Morespecifically, the pump 502 may be a turbopump 502 having a turbinesection 510 and centrifugal pump section 512. While the turbopump 502will be described herein with respect to a centrifugal pump section 512it will be understood that any suitable pump section may be utilized inthe turbopump 502 including a reciprocating pump, a screw pump, a vanepump, a gear pump, etc. The type of pump section utilized may varydepending upon the specific fluid being pumped.

As illustrated, the turbine section 510 may be fluidly coupled to thecompressor section of the turbine engine 506, indicated at 504, suchthat the turbopump 502 may be powered utilizing the pressurized air fromthe compressor discharge pressure of the gas turbine 506. The expansionof pressurized air through the turbine section 510 of the turbopump 502will power the centrifugal pump section 512, effecting its rotation suchthat the centrifugal pump section 512 pressurizes liquid cryogenic fuelin the cryogenic fuel system 500. In this manner, the turbopump 502utilizes mechanical energy from pressurized air bled off the compressordischarge pressure. This mechanical energy in turn powers a centrifugalpump section 512 connected to the same shaft as the air turbine in theturbine section 510 to pressurize liquid cryogenic fuels. The cryogenicfuel must be at a high enough pressure to overcome pressure drop acrossall components in the line from the turbopump 502 all the way to thefuel nozzle(s) 518 as well as the pressure delta across the fuelnozzle(s) 518 into the combustor.

In order to ensure the centrifugal pump section 512 does not cavitate, asmall pump 520 may be installed upstream of the turbopump 502 toincrease fuel pressure and provide margin to fluid saturation point. Thepump 520 may fluidly couple the tank 515 to the centrifugal pump section512 of the turbopump 502 to increase pressure of the cryogenic fueldelivered to the centrifugal pump section 512 of the turbopump 502. Thepump 520 may be a small electrically driven pressurizing pump or chargepump. The pump 520 could alternatively be powered by other mechanicalmeans instead of an electric motor. Alternatively, other means could beutilized to avoid cavitation including increasing tank pressure orcryo-fuel head on the suction line, cooling the fuel outside of the tankor operating the tank below saturation point.

An air valve 522 may be fluidly coupled between the compressor sectionand the turbopump 502. The flow of air from the compressor, as indicatedat 504, to the turbopump 502 may be modulated utilizing the air valve522; the modulation of air to the turbopump 502 in turn may control thevolume flow rate of the cryogenic fuel.

It will be understood that any suitable turbopump 502 may be utilized.An alternative turbopump 502 may utilize staged combustion of the fuelwith compressor discharge pressure air or a full gas generator conceptin lieu of the turbine expander section of the pump. Using stagedcombustion or a full gas generator may increase complexity of theturbine section but could lead to a reduction in the amount ofcompressor discharge pressure air required to operate the turbopump 502at full flow conditions. Additionally, an electric motor/generator couldbe incorporated on the same shaft as the turbopump 502 to either extractelectric power from the turbopump 502 or use electric power to reducethe amount of compressor discharge pressure air extracted from the gasturbine compressor. Instead of a rotary turbopump, either the pump orthe turbine section or both could be replaced by reciprocating pistonsor other mechanical means to extract energy from pressurized fluid orimpart pressure onto fluid.

The exemplary system above may be used in a method for deliveringcryogenic fuel in a cryogenic fuel system to a jet engine having acombustion section and a compressor section. The method may includepressurizing the cryogenic fuel provided to the combustion section ofthe jet engine utilizing a turbopump driven by compressor dischargepressure from the jet engine. The cryogenic fuel provided to thecombustion section may be pressurized by the turbopump to a pressurehigher than the pressure of air coming into the combustion section ofthe jet engine from the compressor section of the jet engine. Further,the cryogenic fuel may be pre-pressurized prior to the cryogenic fuelbeing pressurized by the turbopump. Further still, the flow of thecompressor discharge from the jet engine to the turbopump may bemodulated. Further still, it will be understood that the application offuel pressurization using a turbopump and high pressure compressor bleedfrom a gas turbine is not limited to LNG type fuels. Such a system andmethod may also be used with other types of fuel such as Jet A, naturalgas, bio-fuels, etc.

The above described embodiment provides a variety of benefits includingthat utilizing an air turbine to power the centrifugal pump sectionpermits operation of the pump at high rotational speeds to achieveoptimum pump efficiency. Utilizing turbopump technology as a means toincrease cryo-fuel pressure in lieu of a pressurized tank or an electricpump presents significant technical and commercial advantages. Comparedto a pressurized tank, the turbopump is less complex, reduces safetyrisks involved with high pressure fuel tanks, reduces tank heaterrequirements, and greatly reduces tank weight, because less structuralreinforcement is necessary at low pressures. Compared to an electricallydriven pump, the turbopump does not need an external electric powersource; the pump can be operated at higher speeds for optimumefficiency, and allows for fuel flow rate modulation without the needfor complex power electronics. In addition, the turbopump weighssignificantly less and occupies less space. The compact size and weightof the turbopump allows for installation of the turbopump on the engine,reducing the length of pressurized fuel lines which need to be runbetween the pump and engine nozzles. Shorter pressurized fuel linespresent a safety as well as weight advantage.

The propulsion system 100 comprises a gas turbine engine 101 thatgenerates the propulsive thrust by burning a fuel in a combustor. FIG. 5is a schematic view of an exemplary gas turbine engine 101 including afan 103 and a core engine 108 having a high pressure compressor 105, anda combustor 90. Engine 101 also includes a high pressure turbine 155, alow pressure turbine 157, and a booster 104. The exemplary gas turbineengine 101 has a fan 103 that produces at least a portion of thepropulsive thrust. Engine 101 has an intake side 109 and an exhaust side110. Fan 103 and turbine 157 are coupled together using a first rotorshaft 114, and compressor 105 and turbine 155 are coupled together usinga second rotor shaft 115. In some applications, such as, for example,shown in FIG. 5, the fan 103 blade assemblies are at least partiallypositioned within an engine casing 116. In other applications, the fan103 may form a portion of an “open rotor” where there is no casingsurrounding the fan blade assembly.

During operation, air flows axially through fan 103, in a direction thatis substantially parallel to a central line axis 15 extending throughengine 101, and compressed air is supplied to high pressure compressor105. The highly compressed air is delivered to combustor 90. Hot gases(not shown in FIG. 5) from combustor 90 drives turbines 155 and 157.Turbine 157 drives fan 103 by way of shaft 114 and similarly, turbine155 drives compressor 105 by way of shaft 115. In alternativeembodiments, the engine 101 may have an additional compressor, sometimesknown in the art as an intermediate pressure compressor, driven byanother turbine stage (not shown in FIG. 5).

During operation of the aircraft system 5 (See exemplary flight profileshown in FIG. 8), the gas turbine engine 101 in the propulsion system100 may use, for example, the first fuel 11 during a first selectedportion of operation of propulsion system, such as for example, duringtake off. The propulsion system 100 may use the second fuel 12, such as,for example, LNG, during a second selected portion of operation ofpropulsion system such as during cruise. Alternatively, during selectedportions of the operation of the aircraft system 5, the gas turbineengine 101 is capable of generating the propulsive thrust using both thefirst fuel 11 and the second fuel 12 simultaneously. The proportion ofthe first fuel and second fuel may be varied between 0% to 100% asappropriate during various stages of the operation of the propulsionsystem.

An aircraft and engine system, described herein, is capable of operationusing two fuels, one of which may be a cryogenic fuel such as forexample, LNG (liquefied natural gas), the other a conventional kerosenebased jet fuel such as Jet-A, JP-8, JP-5 or similar grades availableworldwide.

The Jet-A fuel system is similar to conventional aircraft fuel systems,with the exception of the fuel nozzles, which are capable of firingJet-A and cryogenic/LNG to the combustor in proportions from 0-100%. Inthe embodiment shown in FIG. 1, the LNG system includes a fuel tank,which optionally contains the following features: (i) vent lines withappropriate check valves to maintain a specified pressure in the tank;(ii) drain lines for the liquid cryogenic fuel; (iii) gauging or othermeasurement capability to assess the temperature, pressure, and volumeof cryogenic (LNG) fuel present in the tank; (iv) a boost pump locatedin the cryogenic (LNG) tank or optionally outside of the tank, whichincreases the pressure of the cryogenic (LNG) fuel to transport it tothe engine; and (iv) an optional cryo-cooler to keep the tank atcryogenic temperatures indefinitely.

The fuel tank will preferably operate at or near atmospheric pressure,but can operate in the range of 0 to 100 psig. Alternative embodimentsof the fuel system may include high tank pressures and temperatures. Thecryogenic (LNG) fuel lines running from the tank and boost pump to theengine pylons may have the following features: (i) single or double wallconstruction; (ii) vacuum insulation or low thermal conductivitymaterial insulation; and (iii) an optional cryo-cooler to re-circulateLNG flow to the tank without adding heat to the LNG tank. The cryogenic(LNG) fuel tank can be located in the aircraft where a conventionalJet-A auxiliary fuel tank is located on existing systems, for example,in the forward or aft cargo hold. Alternatively, a cryogenic (LNG) fueltank can be located in the center wing tank location. An auxiliary fueltank utilizing cryogenic (LNG) fuel may be designed so that it can beremoved if cryogenic (LNG) fuel will not be used for an extended periodof time.

A high pressure pump may be located in the pylon or on board the engineto raise the pressure of the cryogenic (LNG) fuel to levels sufficientto inject fuel into the gas turbine combustor. The pump may or may notraise the pressure of the LNG/cryogenic liquid above the criticalpressure (Pc) of cryogenic (LNG) fuel. A heat exchanger, referred toherein as a “vaporizer,” which may be mounted on or near the engine,adds thermal energy to the liquefied natural gas fuel, raising thetemperature and volumetrically expanding the cryogenic (LNG) fuel. Heat(thermal energy) from the vaporizer can come from many sources. Theseinclude, but are not limited to: (i) the gas turbine exhaust; (ii)compressor intercooling; (iii) high pressure and/or low pressure turbineclearance control air; (iv) LPT pipe cooling parasitic air; (v) cooledcooling air from the HP turbine; (vi) lubricating oil; or (vii) on boardavionics or electronics. The heat exchanger can be of various designs,including shell and tube, double pipe, fin plate, etc., and can flow ina co-current, counter current, or cross current manner. Heat exchangecan occur in direct or indirect contact with the heat sources listedabove.

A control valve is located downstream of the vaporizer/heat exchangeunit described above. The purpose of the control valve is to meter theflow to a specified level into the fuel manifold across the range ofoperational conditions associated with the gas turbine engine operation.A secondary purpose of the control valve is to act as a back pressureregulator, setting the pressure of the system above the criticalpressure of cryogenic (LNG) fuel.

A fuel manifold is located downstream of the control valve, which servesto uniformly distribute gaseous fuel to the gas turbine fuel nozzles. Insome embodiments, the manifold can optionally act as a heat exchanger,transferring thermal energy from the core cowl compartment or otherthermal surroundings to the cryogenic/LNG/natural gas fuel. A purgemanifold system can optionally be employed with the fuel manifold topurge the fuel manifold with compressor air (CDP) when the gaseous fuelsystem is not in operation. This will prevent hot gas ingestion into thegaseous fuel nozzles due to circumferential pressure variations.Optionally, check valves in or near the fuel nozzles can prevent hot gasingestion.

An exemplary embodiment of the system described herein may operate asfollows: Cryogenic (LNG) fuel is located in the tank at about 15 psiaand about −265° F. It is pumped to approximately 30 psi by the boostpump located on the aircraft. Liquid cryogenic (LNG) fuel flows acrossthe wing via insulated double walled piping to the aircraft pylon whereit is stepped up to about 100 to 1,500 psia and can be above or belowthe critical pressure of natural gas/methane. The cryogenic (LNG) fuelis then routed to the vaporizer where it volumetrically expands to agas. The vaporizer may be sized to keep the Mach number andcorresponding pressure losses low. Gaseous natural gas is then meteredthough a control valve and into the fuel manifold and fuel nozzles whereit is combusted in an otherwise standard aviation gas turbine enginesystem, providing thrust to the airplane. As cycle conditions change,the pressure in the boost pump (about 30 psi for example) and thepressure in the HP pump (about 1,000 psi for example) are maintained atan approximately constant level. Flow is controlled by the meteringvalve. The variation in flow in combination with the appropriately sizedfuel nozzles result in acceptable and varying pressures in the manifold.

The exemplary aircraft system 5 has a fuel delivery system fordelivering one or more types of fuels from the storage system 10 for usein the propulsion system 100. For a conventional liquid fuel such as,for example, a kerosene based jet fuel, a conventional fuel deliverysystem may be used. The exemplary fuel delivery system described herein,and shown schematically in FIGS. 3 and 4, comprises a cryogenic fueldelivery system 50 for an aircraft system 5. The exemplary fuel system50 shown in FIG. 3 comprises a cryogenic fuel tank 122 capable ofstoring a cryogenic liquid fuel 112. In one embodiment, the cryogenicliquid fuel 112 is LNG. Other alternative cryogenic liquid fuels mayalso be used. In the exemplary fuel system 50, the cryogenic liquid fuel112, such as, for example, LNG, is at a first pressure “P1”. Thepressure P1 is preferably close to atmospheric pressure, such as, forexample, 15 psia.

The exemplary fuel system 50 has a boost pump 52 such that it is in flowcommunication with the cryogenic fuel tank 122. During operation, whencryogenic fuel is needed in the dual fuel propulsion system 100, theboost pump 52 removes a portion of the cryogenic liquid fuel 112 fromthe cryogenic fuel tank 122 and increases its pressure to a secondpressure “P2” and flows it into a wing supply conduit 54 located in awing 7 of the aircraft system 5. The pressure P2 is chosen such that theliquid cryogenic fuel maintains its liquid state (L) during the flow inthe supply conduit 54. The pressure P2 may be in the range of about 30psia to about 40 psia. Based on analysis using known methods, for LNG,30 psia is found to be adequate. The boost pump 52 may be located at asuitable location in the fuselage 6 of the aircraft system 5.Alternatively, the boost pump 52 may be located close to the cryogenicfuel tank 122. In other embodiments, the boost pump 52 may be locatedinside the cryogenic fuel tank 122. In order to substantially maintain aliquid state of the cryogenic fuel during delivery, at least a portionof the wing supply conduit 54 is insulated. In some exemplaryembodiments, at least a portion of the conduit 54 has a double wallconstruction. The conduits 54 and the boost pump 52 may be made usingknown materials such as titanium, Inconel, aluminum or compositematerials.

The exemplary fuel system 50 has a high-pressure pump 58 that is in flowcommunication with the wing supply conduit 54 and is capable ofreceiving the cryogenic liquid fuel 112 supplied by the boost pump 52.The high-pressure pump 58 increases the pressure of the liquid cryogenicfuel (such as, for example, LNG) to a third pressure “P3” sufficient toinject the fuel into the propulsion system 100. The pressure P3 may bein the range of about 100 psia to about 1000 psia. The high-pressurepump 58 may be located at a suitable location in the aircraft system 5or the propulsion system 100. The high-pressure pump 58 is preferablylocated in a pylon 55 of aircraft system 5 that supports the propulsionsystem 100.

As shown in FIG. 3, the exemplary fuel system 50 has a vaporizer 60 forchanging the cryogenic liquid fuel 112 into a gaseous (G) fuel 13. Thevaporizer 60 receives the high pressure cryogenic liquid fuel and addsheat (thermal energy) to the cryogenic liquid fuel (such as, forexample, LNG) raising its temperature and volumetrically expanding it.Heat (thermal energy) can be supplied from one or more sources in thepropulsion system 100. For example, heat for vaporizing the cryogenicliquid fuel in the vaporizer may be supplied from one or more of severalsources, such as, for example, the gas turbine exhaust 99, compressor105, high pressure turbine 155, low pressure turbine 157, fan bypass107, turbine cooling air, lubricating oil in the engine, aircraft systemavionics/electronics, or any source of heat in the propulsion system100. Due to the exchange of heat that occurs in the vaporizer 60, thevaporizer 60 may be alternatively referred to as a heat exchanger. Theheat exchanger portion of the vaporizer 60 may include a shell and tubetype heat exchanger, or a double pipe type heat exchanger, orfin-and-plate type heat exchanger. The hot fluid and cold fluid flow inthe vaporizer may be co-current, or counter-current, or a cross currentflow type. The heat exchange between the hot fluid and the cold fluid inthe vaporizer may occur directly through a wall or indirectly, using anintermediate work fluid.

The cryogenic fuel delivery system 50 comprises a flow metering valve 65(“FMV”, also referred to as a Control Valve) that is in flowcommunication with the vaporizer 60 and a manifold 70. The flow meteringvalve 65 is located downstream of the vaporizer/heat exchange unitdescribed above. The purpose of the FMV (control valve) is to meter thefuel flow to a specified level into the fuel manifold 70 across therange of operational conditions associated with the gas turbine engineoperation. A secondary purpose of the control valve is to act as a backpressure regulator, setting the pressure of the system above thecritical pressure of the cryogenic fuel such as LNG. The flow meteringvalve 65 receives the gaseous fuel 13 supplied from the vaporizer andreduces its pressure to a fourth pressure “P4”. The manifold 70 iscapable of receiving the gaseous fuel 13 and distributing it to a fuelnozzle 80 in the gas turbine engine 101. In a preferred embodiment, thevaporizer 60 changes the cryogenic liquid fuel 112 into the gaseous fuel13 at a substantially constant pressure. FIG. 3a schematically shows thestate and pressure of the fuel at various points in the delivery system50.

The cryogenic fuel delivery system 50 further comprises a plurality offuel nozzles 80 located in the gas turbine engine 101. The fuel nozzle80 delivers the gaseous fuel 13 into the combustor 90 for combustion.The fuel manifold 70, located downstream of the control valve 65, servesto uniformly distribute gaseous fuel 13 to the gas turbine fuel nozzles80. In some embodiments, the manifold 70 can optionally act as a heatexchanger, transferring thermal energy from the propulsion system corecowl compartment or other thermal surroundings to the LNG/natural gasfuel. In one embodiment, the fuel nozzle 80 is configured to selectivelyreceive a conventional liquid fuel (such as the conventional kerosenebased liquid fuel) or the gaseous fuel 13 generated by the vaporizerfrom the cryogenic liquid fuel such as LNG. In another embodiment, thefuel nozzle 80 is configured to selectively receive a liquid fuel andthe gaseous fuel 13 and configured to supply the gaseous fuel 13 and aliquid fuel to the combustor 90 to facilitate co-combustion of the twotypes of fuels. In another embodiment, the gas turbine engine 101comprises a plurality of fuel nozzles 80 wherein some of the fuelnozzles 80 are configured to receive a liquid fuel and some of the fuelnozzles 80 are configured to receive the gaseous fuel 13 and arrangedsuitably for combustion in the combustor 90.

In another embodiment of the present invention, fuel manifold 70 in thegas turbine engine 101 comprises an optional purge manifold system topurge the fuel manifold with compressor air, or other air, from theengine when the gaseous fuel system is not in operation. This willprevent hot gas ingestion into the gaseous fuel nozzles due tocircumferential pressure variations in the combustor 90. Optionally,check valves in or near the fuel nozzles can be used prevent hot gasingestion in the fuel nozzles or manifold.

In an exemplary dual fuel gas turbine propulsion system described hereinthat uses LNG as the cryogenic liquid fuel is described as follows: LNGis located in the tank 22, 122 at 15 psia and −265° F. It is pumped toapproximately 30 psi by the boost pump 52 located on the aircraft.Liquid LNG flows across the wing 7 via insulated double walled piping 54to the aircraft pylon 55 where it is stepped up to 100 to 1,500 psia andmay be above or below the critical pressure of natural gas/methane. TheLiquefied Natural Gas is then routed to the vaporizer 60 where itvolumetrically expands to a gas. The vaporizer 60 is sized to keep theMach number and corresponding pressure losses low. Gaseous natural gasis then metered though a control valve 65 and into the fuel manifold 70and fuel nozzles 80 where it is combusted in an dual fuel aviation gasturbine system 100, 101, providing thrust to the aircraft system 5. Ascycle conditions change, the pressure in the boost pump (30 psi) and thepressure in the HP pump 58 (1,000 psi) are maintained at anapproximately constant level. Flow is controlled by the metering valve65. The variation in flow in combination with the appropriately sizedfuel nozzles result in acceptable and varying pressures in the manifold.

The dual fuel system consists of parallel fuel delivery systems forkerosene based fuel (Jet-A, JP-8, JP-5, etc) and a cryogenic fuel (LNGfor example). The kerosene fuel delivery is substantially unchanged fromthe current design, with the exception of the combustor fuel nozzles,which are designed to co-fire kerosene and natural gas in anyproportion. As shown in FIG. 3, the cryogenic fuel (LNG for example)fuel delivery system consists of the following features: (A) A dual fuelnozzle and combustion system, capable of utilizing cryogenic fuel (LNGfor example), and Jet-A in any proportion from 0- to 100%; (B) A fuelmanifold and delivery system that also acts as a heat exchanger, heatingcryogenic fuel (LNG for example) to a gas or a supercritical fluid. Themanifold system is designed to concurrently deliver fuel to thecombustor fuel nozzles in a uniform manner, and absorb heat from thesurrounding core cowl, exhaust system, or other heat source, eliminatingor minimizing the need for a separate heat exchanger; (C) A fuel systemthat pumps up cryogenic fuel (LNG for example) in its liquid state aboveor below the critical pressure and adds heat from any of a number ofsources; (D) A low pressure cryo-pump submerged in the cryogenic fuel(LNG for example) fuel tank (optionally located outside the fuel tank);(E) A high pressure cryo-pump located in the aircraft pylon oroptionally on board the engine or nacelle to pump to pressures above thecritical pressure of cryogenic fuel (LNG for example). (F) A purgemanifold system can optionally employed with the fuel manifold to purgethe fuel manifold with compressor CDP air when the gaseous fuel systemis not in operation. This will prevent hot gas ingestion into thegaseous fuel nozzles due to circumferential pressure variations.Optionally, check valves in or near the fuel nozzles can prevent hot gasingestion. (G) cryogenic fuel (LNG for example) lines running from thetank and boost pump to the engine pylons have the following features:(1) Single or double wall construction. (2) Vacuum insulation oroptionally low thermal conductivity insulation material such asaerogels. (3) An optional cryo-cooler to recirculate cryogenic fuel (LNGfor example) flow to the tank without adding heat to the cryogenic fuel(LNG for example) tank. (H) A high pressure pump located in the pylon oron board the engine. This pump will raise the pressure of the cryogenicfuel (LNG for example) to levels sufficient to inject natural gas fuelinto the gas turbine combustor. The pump may or may not raise thepressure of the cryogenic liquid (LNG for example) above the criticalpressure (Pc) of cryogenic fuel (LNG for example).

III. A Fuel Storage System

The exemplary aircraft system 5 shown in FIG. 1 comprises a cryogenicfuel storage system 10, such as shown for example, in FIG. 4, forstoring a cryogenic fuel. The exemplary cryogenic fuel storage system 10comprises a cryogenic fuel tank 22, 122 having a first wall 23 forming astorage volume 24 capable of storing a cryogenic liquid fuel 12 such asfor example LNG. As shown schematically in FIG. 4, the exemplarycryogenic fuel storage system 10 has an inflow system 32 capable offlowing the cryogenic liquid fuel 12 into the storage volume 24 and anoutflow system 30 adapted to deliver the cryogenic liquid fuel 12 fromthe cryogenic fuel storage system 10. It further comprises a vent system40 capable of removing at least a portion of a gaseous fuel 19 (that maybe formed during storage) from the cryogenic liquid fuel 12 in thestorage volume 24.

The exemplary cryogenic fuel storage system 10 shown in FIG. 4 furthercomprises a recycle system 34 that is adapted to return at least aportion 29 of unused gaseous fuel 19 into the cryogenic fuel tank 22. Inone embodiment, the recycle system 34 comprises a cryo-cooler 42 thatcools the portion 29 of unused gaseous fuel 19 prior to returning itinto the cryogenic fuel tank 22, 122. An exemplary operation of thecryo-cooler 42 operation is as follows: In an exemplary embodiment, boiloff from the fuel tank can be re-cooled using a reverse Rankinerefrigeration system, also known as a cryo-cooler. The cryo-cooler canbe powered by electric power coming from any of the available systems onboard the aircraft system 5, or, by ground based power systems such asthose which may be available while parked at a boarding gate. Thecryo-cooler system can also be used to re-liquefy natural gas in thefuel system during the dual fuel aircraft gas turbine engine 101 co-firetransitions.

The fuel storage system 10 may further comprise a safety release system45 adapted to vent any high pressure gases that may be formed in thecryogenic fuel tank 22. In one exemplary embodiment, shown schematicallyin FIG. 4, the safety release system 45 comprises a rupture disk 46 thatforms a portion of the first wall 23. The rupture disk 46 is a safetyfeature, designed using known methods, to blow out and release any highpressure gases in the event of an over pressure inside the fuel tank 22.

The cryogenic fuel tank 22 may have a single wall construction or amultiple wall construction. For example, the cryogenic fuel tank 22 mayfurther comprise (See FIG. 4 for example) a second wall 25 thatsubstantially encloses the first wall 23. In one embodiment of the tank,there is a gap 26 between the first wall 23 and the second wall 25 inorder to thermally insulate the tank to reduce heat flow across the tankwalls. In one exemplary embodiment, there is a vacuum in the gap 26between the first wall 23 and the second wall 25. The vacuum may becreated and maintained by a vacuum pump 28. Alternatively, in order toprovide thermal insulation for the tank, the gap 26 between the firstwall 23 and the second wall 25 may be substantially filled with a knownthermal insulation material 27, such as, for example, Aerogel. Othersuitable thermal insulation materials may be used. Baffles 17 may beincluded to control movement of liquid within the tank.

The cryogenic fuel storage system 10 shown in FIG. 4 comprises theoutflow system 30 having a delivery pump 31. The delivery pump may belocated at a convenient location near the tank 22. In order to reduceheat transfer in to the cryogenic fuel, it may be preferable to locatethe delivery pump 31 in the cryogenic fuel tank 22 as shownschematically in FIG. 4. The vent system 40 vents any gases that may beformed in the fuel tank 22. These vented gases may be utilized inseveral useful ways in the aircraft system 5. A few of these are shownschematically in FIG. 4. For example at least a portion of the gaseousfuel 19 may be supplied to the aircraft propulsion system 100 forcooling or combustion in the engine. In another embodiment, the ventsystem 40 supplies at least a portion of the gaseous fuel 19 to a burnerand further venting the combustion products from the burner safelyoutside the aircraft system 5. In another embodiment the vent system 40supplies at least a portion of the gaseous fuel 19 to an auxiliary powerunit 180 that supplies auxiliary power to the aircraft system 5. Inanother embodiment the vent system 40 supplies at least a portion of thegaseous fuel 19 to a fuel cell 182 that produces power. In anotherembodiment the vent system 40 releases at least a portion of the gaseousfuel 19 outside the cryogenic fuel tank 22.

The exemplary operation of the fuel storage system, its componentsincluding the fuel tank, and exemplary sub systems and components isdescribed as follows.

Natural gas exists in liquid form (LNG) at temperatures of approximatelyabout −260° F. and atmospheric pressure. To maintain these temperaturesand pressures on board a passenger, cargo, military, or general aviationaircraft, the features identified below, in selected combinations, allowfor safe, efficient, and cost effective storage of LNG. Referring toFIG. 4, these include:

(A) A fuel tank 21, 22 constructed of alloys such as, but not limitedto, aluminum AL 5456 and higher strength aluminum AL 5086 or othersuitable alloys.

(B) A fuel tank 21, 22 constructed of light weight composite material.

(C) The above tanks 21, 22 with a double wall vacuum feature forimproved insulation and greatly reduced heat flow to the LNG fluid. Thedouble walled tank also acts as a safety containment device in the rarecase where the primary tank is ruptured.

(D) An alternative embodiment of either the above utilizing lightweightinsulation 27, such as, for example, Aerogel, to minimize heat flow fromthe surroundings to the LNG tank and its contents. Aerogel insulationcan be used in addition to, or in place of a double walled tank design.

(E) An optional vacuum pump 28 designed for active evacuation of thespace between the double walled tank. The pump can operate off of LNGboil off fuel, LNG, Jet-A, electric power or any other power sourceavailable to the aircraft.

(F) An LNG tank with a cryogenic pump 31 submerged inside the primarytank for reduced heat transfer to the LNG fluid.

(G) An LNG tank with one or more drain lines 36 capable of removing LNGfrom the tank under normal or emergency conditions. The LNG drain line36 is connected to a suitable cryogenic pump to increase the rate ofremoval beyond the drainage rate due to the LNG gravitational head.

(H) An LNG tank with one or more vent lines 41 for removal of gaseousnatural gas, formed by the absorption of heat from the externalenvironment. This vent line 41 system maintains the tank at a desiredpressure by the use of a 1 way relief valve or back pressure valve 39.

(I) An LNG tank with a parallel safety relief system 45 to the main ventline, should an overpressure situation occur. A burst disk is analternative feature or a parallel feature 46. The relief vent woulddirect gaseous fuel overboard.

(J) An LNG fuel tank, with some or all of the design features above,whose geometry is designed to conform to the existing envelopeassociated with a standard Jet-A auxiliary fuel tank such as thosedesigned and available on commercially available aircrafts.

(K) An LNG fuel tank, with some or all of the design features above,whose geometry is designed to conform to and fit within the lower cargohold(s) of conventional passenger and cargo aircraft such as those foundon commercially available aircrafts.

(L) Modifications to the center wing tank 22 of an existing or newaircraft to properly insulate the LNG, tank, and structural elements.

Venting and boil off systems are designed using known methods. Boil offof LNG is an evaporation process which absorbs energy and cools the tankand its contents. Boil off LNG can be utilized and/or consumed by avariety of different processes, in some cases providing useful work tothe aircraft system, in other cases, simply combusting the fuel for amore environmentally acceptable design. For example, vent gas from theLNG tank consists primarily of methane and is used for any or allcombinations of the following:

(A) Routing to the Aircraft APU (Auxiliary Power Unit) 180. As shown inFIG. 4, a gaseous vent line from the tank is routed in series or inparallel to an Auxiliary Power Unit for use in the combustor. The APUcan be an existing APU, typically found aboard commercial and militaryaircraft, or a separate APU dedicated to converting natural gas boil offto useful electric and/or mechanical power. A boil off natural gascompressor is utilized to compress the natural gas to the appropriatepressure required for utilization in the APU. The APU, in turn, provideselectric power to any system on the engine or A/C.

(B) Routing to one or more aircraft gas turbine engine(s) 101. As shownin FIG. 4, a natural gas vent line from the LNG fuel tank is routed toone or more of the main gas turbine engines 101 and provides anadditional fuel source to the engine during operation. A natural gascompressor is utilized to pump the vent gas to the appropriate pressurerequired for utilization in the aircraft gas turbine engine.

(C) Flared. As shown in FIG. 4, a natural gas vent line from the tank isrouted to a small, dedicated vent combustor 190 with its own electricspark ignition system. In this manner methane gas is not released to theatmosphere. The products of combustion are vented, which results in amore environmentally acceptable system.

(D) Vented. As shown in FIG. 4, a natural gas vent line from the tank isrouted to the exhaust duct of one or more of the aircraft gas turbines.Alternatively, the vent line can be routed to the APU exhaust duct or aseparate dedicated line to any of the aircraft trailing edges. Naturalgas may be suitably vented to atmosphere at one or more of theselocations V.

(E) Ground operation. As shown in FIG. 4, during ground operation, anyof the systems can be designed such that a vent line 41 is attached toground support equipment, which collects and utilizes the natural gasboil off in any ground based system. Venting can also take place duringrefueling operations with ground support equipment that cansimultaneously inject fuel into the aircraft LNG tank using an inflowsystem 32 and capture and reuse vent gases (simultaneous venting andfueling indicated as (S) in FIG. 4).

IV. Propulsion (Engine) System

FIG. 5 shows an exemplary dual fuel propulsion system 100 comprising agas turbine engine 101 capable of generating a propulsive thrust using acryogenic liquid fuel 112. The gas turbine engine 101 comprises acompressor 105 driven by a high-pressure turbine 155 and a combustor 90that burns a fuel and generates hot gases that drive the high-pressureturbine 155. The combustor 90 is capable of burning a conventionalliquid fuel such as kerosene based fuel. The combustor 90 is alsocapable of burning a cryogenic fuel, such as, for example, LNG, that hasbeen suitably prepared for combustion, such as, for example, by avaporizer 60. FIG. 5 shows schematically a vaporizer 60 capable ofchanging the cryogenic liquid fuel 112 into a gaseous fuel 13. The dualfuel propulsion system 100 gas turbine engine 101 further comprises afuel nozzle 80 that supplies the gaseous fuel 13 to the combustor 90 forignition. In one exemplary embodiment, the cryogenic liquid fuel 112used is Liquefied Natural Gas (LNG). In a turbo-fan type dual fuelpropulsion system 100 (shown in FIG. 5 for example) the gas turbineengine 101 comprises a fan 103 located axially forward from thehigh-pressure compressor 105. A booster 104 (shown in FIG. 5) may belocated axially between the fan 103 and the high-pressure compressor 105wherein the fan and booster are driven by a low-pressure turbine 157. Inother embodiments, the dual fuel propulsion system 100 gas turbineengine 101 may include an intermediate pressure compressor driven by anintermediate pressure turbine (both not shown in FIG. 5). The booster104 (or an intermediate pressure compressor) increases the pressure ofthe air that enters the compressor 105 and facilitates the generation ofhigher pressure ratios by the compressor 105. In the exemplaryembodiment shown in FIG. 5, the fan and the booster are driven by thelow pressure turbine 157, and the high pressure compressor is driven thehigh pressure turbine 155.

The vaporizer 60, shown schematically in FIG. 5, is mounted on or nearthe engine 101. One of the functions of the vaporizer 60 is to addthermal energy to the cryogenic fuel, such as the liquefied natural gas(LNG) fuel, raising its temperature. In this context, the vaporizerfunctions as heat exchanger. Another, function of the vaporizer 60 is tovolumetrically expand the cryogenic fuel, such as the liquefied naturalgas (LNG) fuel to a gaseous form for later combustion. Heat (thermalenergy) for use in the vaporizer 60 can come from or more of manysources in the propulsion system 100 and aircraft system 5. Theseinclude, but are not limited to: (i) The gas turbine exhaust, (ii)Compressor intercooling, (iii) High pressure and/or low pressure turbineclearance control air, (iv) LPT pipe cooling parasitic air, (v) coolingair used in the High pressure and/or low pressure turbine, (vi)Lubricating oil, and (vii) On board avionics, electronics in theaircraft system 5. The heat for the vaporizer may also be supplied fromthe compressor 105, booster 104, intermediate pressure compressor (notshown) and/or the fan bypass air stream 107 (See FIG. 5). An exemplaryembodiment using a portion of the discharge air from the compressor 105is shown in FIG. 6. A portion of the compressor discharge air 2 is bledout to the vaporizer 60, as shown by item 3 in FIG. 6. The cryogenicliquid fuel 21, such as for example, LNG, enters vaporizer 60 whereinthe heat from the airflow stream 3 is transferred to the cryogenicliquid fuel 21. In one exemplary embodiment, the heated cryogenic fuelis further expanded, as described previously herein, producing gaseousfuel 13 in the vaporizer 60. The gaseous fuel 13 is then introduced intocombustor 90 using a fuel nozzle 80 (See FIG. 6). The cooled airflow 4that exits from the vaporizer can be used for cooling other enginecomponents, such as the combustor 90 structures and/or the high-pressureturbine 155 structures. The heat exchanger portion in the vaporizer 60can be of a known design, such as for example, shell and tube design,double pipe design, and/or fin plate design. The fuel 112 flow directionand the heating fluid 96 direction in the vaporizer 60 (see FIG. 5) maybe in a co-current direction, counter-current direction, or they mayflow in a cross-current manner to promote efficient heat exchangebetween the cryogenic fuel and the heating fluid.

Heat exchange in the vaporizer 60 can occur in direct manner between thecryogenic fuel and the heating fluid, through a metallic wall. FIG. 6shows schematically a direct heat exchanger in the vaporizer 60. FIG. 7ashows schematically an exemplary direct heat exchanger 63 that uses aportion 97 of the gas turbine engine 101 exhaust gas 99 to heat thecryogenic liquid fuel 112. Alternatively, heat exchange in the vaporizer60 can occur in an indirect manner between the cryogenic fuel and theheat sources listed above, through the use of an intermediate heatingfluid. FIG. 7b shows an exemplary vaporizer 60 that uses an indirectheat exchanger 64 that uses an intermediary heating fluid 68 to heat thecryogenic liquid fuel 112. In such an indirect heat exchanger shown inFIG. 7b , the intermediary heating fluid 68 is heated by a portion 97 ofthe exhaust gas 99 from the gas turbine engine 101. Heat from theintermediary heating fluid 68 is then transferred to the cryogenicliquid fuel 112. FIG. 7c shows another embodiment of an indirectexchanger used in a vaporizer 60. In this alternative embodiment, theintermediary heating fluid 68 is heated by a portion of a fan bypassstream 107 of the gas turbine engine 101, as well as a portion 97 of theengine exhaust gas 99. The intermediary heating fluid 68 then heats thecryogenic liquid fuel 112. A control valve 38 is used to control therelative heat exchanges between the flow streams.

(V) Method of Operating Dual Fuel Aircraft System

An exemplary method of operation of the aircraft system 5 using a dualfuel propulsion system 100 is described as follows with respect to anexemplary flight mission profile shown schematically in FIG. 8. Theexemplary flight mission profile shown schematically in FIG. 8 shows theEngine power setting during various portions of the flight missionidentified by the letter labels A-B-C-D-E- . . . -X-Y etc. For example,A-B represents the start, B-C shows ground-idle, G-H shows take-off, T-Land O-P show cruise, etc. During operation of the aircraft system 5 (Seeexemplary flight profile 120 in FIG. 8), the gas turbine engine 101 inthe propulsion system 100 may use, for example, the first fuel 11 duringa first selected portion of operation of propulsion system, such as forexample, during take off. The propulsion system 100 may use the secondfuel 12, such as, for example, LNG, during a second selected portion ofoperation of propulsion system such as during cruise. Alternatively,during selected portions of the operation of the aircraft system 5, thegas turbine engine 101 is capable of generating the propulsive thrustusing both the first fuel 11 and the second fuel 12 simultaneously. Theproportion of the first fuel and second fuel may be varied between 0% to100% as appropriate during various stages of the operation of the dualfuel propulsion system 100.

An exemplary method of operating a dual fuel propulsion system 100 usinga dual fuel gas turbine engine 101 comprises the following steps of:starting the aircraft engine 101 (see A-B in FIG. 8) by burning a firstfuel 11 in a combustor 90 that generates hot gases that drive a gasturbine in the engine 101. The first fuel 11 may be a known type ofliquid fuel, such as a kerosene based Jet Fuel. The engine 101, whenstarted, may produce enough hot gases that may used to vaporize a secondfuel, such as, for example, a cryogenic fuel. A second fuel 12 is thenvaporized using heat in a vaporizer 60 to form a gaseous fuel 13. Thesecond fuel may be a cryogenic liquid fuel 112, such as, for example,LNG. The operation of an exemplary vaporizer 60 has been describedherein previously. The gaseous fuel 13 is then introduced into thecombustor 90 of the engine 101 using a fuel nozzle 80 and the gaseousfuel 13 is burned in the combustor 90 that generates hot gases thatdrive the gas turbine in the engine. The amount of the second fuelintroduced into the combustor may be controlled using a flow meteringvalve 65. The exemplary method may further comprise the step of stoppingthe supply of the first fuel 11 after starting the aircraft engine, ifdesired.

In the exemplary method of operating the dual fuel aircraft gas turbineengine 101, the step of vaporizing the second fuel 12 may be performedusing heat from a hot gas extracted from a heat source in the engine101. As described previously, in one embodiment of the method, the hotgas may be compressed air from a compressor 155 in the engine (forexample, as shown in FIG. 6). In another embodiment of the method, thehot gas is supplied from an exhaust nozzle 98 or exhaust stream 99 ofthe engine (for example, as shown in FIG. 7a ).

The exemplary method of operating a dual fuel aircraft engine 101, may,optionally, comprise the steps of using a selected proportion of thefirst fuel 11 and a second fuel 12 during selected portions of a flightprofile 120, such as shown, for example, in FIG. 8, to generate hotgases that drive a gas turbine engine 101. The second fuel 12 may be acryogenic liquid fuel 112, such as, for example, Liquefied Natural Gas(LNG). In the method above, the step of varying the proportion of thefirst fuel 12 and the second fuel 13 during different portions of theflight profile 120 (see FIG. 8) may be used to advantage to operate theaircraft system in an economic and efficient manner. This is possible,for example, in situations where the cost of the second fuel 12 is lowerthan the cost of the first fuel 11. This may be the case, for example,while using LNG as the second fuel 12 and kerosene based liquid fuelssuch as Jet-A fuel, as first fuel 11. In the exemplary method ofoperating a dual fuel aircraft engine 101, the proportion (ratio) ofamount of the second fuel 12 used to the amount of the first fuel usedmay be varied between about 0% and 100%, depending on the portion of theflight mission. For example, in one exemplary method, the proportion ofa cheaper second fuel used (such as LNG) to the kerosene based fuel usedis about 100% during a cruise part of the flight profile, in order tominimize the cost of fuel. In another exemplary operating method, theproportion of the second fuel is about 50% during a take-off part of theflight profile that requires a much higher thrust level.

The exemplary method of operating a dual fuel aircraft engine 101described above may further comprise the step of controlling the amountsof the first fuel 11 and the second fuel 12 introduced into thecombustor 90 using a control system 130. An exemplary control system 130is shown schematically in FIG. 5. The control system 130 sends a controlsignal 131 (S1) to a control valve 135 to control the amount of thefirst fuel 11 that is introduced to the combustor 90. The control system130 also sends another control signal 132 (S2) to a control valve 65 tocontrol the amount of the second fuel 12 that is introduced to thecombustor 90. The proportion of the first fuel 11 and second fuel 12used can be varied between 0% to 100% by a controller 134 that isprogrammed to vary the proportion as required during different flightsegments of the flight profile 120. The control system 130 may alsoreceive a feed back signal 133, based for example on the fan speed orthe compressor speed or other suitable engine operating parameters. Inone exemplary method, the control system may be a part of the enginecontrol system, such as, for example, a Full Authority DigitalElectronic Control (FADEC) 357. In another exemplary method, amechanical or hydromechanical engine control system may form part or allof the control system.

The control system 130, 357 architecture and strategy is suitablydesigned to accomplish economic operation of the aircraft system 5.Control system feedback to the boost pump 52 and high pressure pump(s)58 can be accomplished via the Engine FADEC 357 or by distributedcomputing with a separate control system that may, optionally,communicate with the Engine FADEC and with the aircraft system 5 controlsystem through various available data busses.

The control system, such as for example, shown in FIG. 5, item 130, mayvary pump 52, 58 speed and output to maintain a specified pressureacross the wing 7 for safety purposes (for example at about 30-40 psi)and a different pressure downstream of the high pressure pump 58 (forexample at about 100 to 1500 psi) to maintain a system pressure abovethe critical point of LNG and avoid two phase flow, and, to reduce thevolume and weight of the LNG fuel delivery system by operation at highpressures and fuel densities.

In an exemplary control system 130, 357, the control system software mayinclude any or all of the following logic: (A) A control system strategythat maximizes the use of the cryogenic fuel such as, for example, LNG,on takeoff and/or other points in the envelope at high compressordischarge temperatures (T3) and/or turbine inlet temperatures (T41); (B)A control system strategy that maximizes the use of cryogenic fuel suchas, for example, LNG, on a mission to minimize fuel costs; (C) A controlsystem 130, 357 that re-lights on the first fuel, such as, for example,Jet-A, only for altitude relights; (D) A control system 130, 357 thatperforms ground starts on conventional Jet-A only as a default setting;(E) A control system 130, 357 that defaults to Jet-A only during any nontypical maneuver; (F) A control system 130, 357 that allows for manual(pilot commanded) selection of conventional fuel (like Jet-A) orcryogenic fuel such as, for example, LNG, in any proportion; (G) Acontrol system 130, 357 that utilizes 100% conventional fuel (likeJet-A) for all fast accels and decels.

To the extent not already described, the different features andstructures of the various embodiments may be used in combination witheach other as desired. That one feature may not be illustrated in all ofthe embodiments is not meant to be construed that it may not be, but isdone for brevity of description. Thus, the various features of thedifferent embodiments may be mixed and matched as desired to form newembodiments, whether or not the new embodiments are expressly described.All combinations or permutations of features described herein arecovered by this disclosure.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A cryogenic fuel system for an aircraft having a turbine engine with a compressor section and a combustion chamber, comprising: a cryogenic fuel stored in a tank; a supply line operably coupling the tank to the combustion chamber; a first pump coupling the tank to the supply line to pump the cryogenic fuel at high pressure through the supply line where the pump is operably coupled to the compressor such that operation of the turbine engine drives the pump; and a second pump fluidly coupling the tank to the first pump to pre-pressurize the cryogenic fuel delivered to the first pump.
 2. The cryogenic fuel system of claim 1, wherein the first pump pressurizes the cryogenic fuel to a pressure higher than the pressure of air coming into the combustion chamber of the turbine engine from the compressor section of the turbine engine.
 3. The cryogenic fuel system of claim 2, wherein the cryogenic fuel is Liquefied Natural Gas (LNG).
 4. The cryogenic fuel system of claim 1, wherein the first pump is a turbopump having a turbine section and a centrifugal pump section.
 5. The cryogenic fuel system of claim 4, wherein the cryogenic fuel is Liquefied Natural Gas (LNG).
 6. The cryogenic fuel system of claim 4, wherein the turbine section of the turbopump is fluidly coupled to the compressor to discharge pressure from the turbine engine such that the discharge pressure from the compressor drives the first pump.
 7. The cryogenic fuel system of claim 6, wherein the turbine section of the turbopump effects rotation of the centrifugal pump section of the turbopump to pressurize the cryogenic fuel in the cryogenic fuel system.
 8. The cryogenic fuel system of claim 7, further comprising an air valve fluidly coupled between the compressor section of the turbine engine and the turbine section of the turbopump and where the air valve is configured to modulate a flow of air to discharge pressure from the compressor to the turbine section of the turbopump.
 9. The cryogenic fuel system of claim 8, further comprising a charge pump fluidly coupling the tank to the centrifugal pump section of the turbopump to increase pressure of the cryogenic fuel delivered to the centrifugal pump section of the turbopump.
 10. The cryogenic fuel system of claim 1, wherein the first pump is a turbopump having a turbine section and a pump section having a pump type selected from one of: a reciprocating pump, a screw pump, a vane pump, or a gear pump.
 11. The cryogenic fuel system of claim 10, wherein the cryogenic fuel is Liquefied Natural Gas (LNG).
 12. The cryogenic fuel system of claim 1, wherein the cryogenic fuel is Liquefied Natural Gas (LNG).
 13. A method for delivering cryogenic fuel in a fuel system to a turbine engine having a combustion section and a compressor section, comprising: pressurizing the cryogenic fuel provided to the combustion section of the turbine engine utilizing a turbopump driven by the compressor to discharge pressure from the turbine engine, and pressurizing the cryogenic fuel with a second fuel pump before said cryogenic fuel is pressurized by the turbopump.
 14. The method of claim 13, wherein the cryogenic fuel provided to the combustion section is pressurized by the turbopump to a pressure higher than the pressure of air coming into the combustion section of the turbine engine from the compressor section of the turbine engine.
 15. The method of claim 14, further comprising modulating the flow of the compressor discharge from the turbine engine to the turbopump.
 16. The method of claim 13, further comprising pre-pressurizing the cryogenic fuel prior to the cryogenic fuel being pressurized by the turbopump.
 17. The method of claim 16, further comprising modulating the flow of the compressor discharge from the turbine engine to the turbopump.
 18. The method of claim 13, further comprising modulating the flow of the compressor discharge from the turbine engine to the turbopump.
 19. The method of claim 13, further comprising modulating the flow of the compressor discharge from the turbine engine to the turbopump. 